1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to an industrial turbine stator vane with an impingement cooling insert for cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A first stage turbine stator vane with an insert for producing impingement cooling is shown in FIG. 1. The vane includes an airfoil 10 with ribs extending across the pressure side wall to the suction side wall to form in this case three impingement cooling air cavities 11, 12 and 13. Each impingement cavity includes an impingement cooling insert with an arrangement of impingement cooling holes 17 directed to discharge impingement cooling air to the backside surfaces of the airfoil walls. A leading edge insert 14 is secured in the leading edge cavity 11, a mid-chord insert 15 is secured within the mid-chord cavity 12, and a trailing edge region insert 16 is secured within the third or trailing edge cavity 13. A showerhead arrangement of film cooling holes 19 are located in the leading edge region of the airfoil, a row of trailing edge exit holes 18 in the trailing edge, and rows of film cooling holes 21 are located on the pressure and suction side walls to discharge film cooling air.
The vane cooling circuit of FIG. 1 works like this. Cooling air supplied to the vane flows into the three cooling air cavities 11-13 and then through the impingement holes 17 formed in the inserts 14-16 to produce impingement cooling on the backside surface of the airfoil walls. The spent impingement cooling air then flows through the film cooling holes spaced around the airfoil surface or out through the exit holes in the trailing edge of the airfoil.